The present invention relates to gas turbine engines and, more particularly, to aircraft type high bypass ratio turbine engines having multi-stage compressor and turbine sections.
A typical modern gas turbine aircraft engine, particularly of the high bypass ratio type, includes multi-stage high pressure compressor and turbine sections interconnected by a central compression shaft or, in some models, a forward shaft. In the latter instance, the forward shaft extends between the webs of the last stage high pressure compressor disk and the first stage high pressure turbine disk webs. The high pressure turbine section typically includes first and second stage disks in which the second stage disk is attached to the first stage disk by a bolted connection. The interstage volume between the first and second stage disks is enclosed by a seal extending between the outer peripheries of the turbine disks. The seal is generally cylindrical in shape and its wall is outwardly convex in shape.
The first and second stage disks are isolated by a forward face plate, attached to the forward face of the first stage disk, and an aft seal attached to the rearward face of the second stage disk web. Typically, cooling air ducted externally from the compressor section is circulated within the volumes defined by the face plate and aft seal, as well as the interstage volume, in order to cool the disks and blades they support. The cooling air is conveyed radially outwardly from the turbine section through channels formed in the turbine blades.
In such engines, virtually all of the connections between components are accomplished through bolting. That is, the forward face plate is connected to the stage one disk by a circular pattern of bolts, extending about the face plate and disk. The inner periphery of the face plate is bolted to a disk positioned forwardly of the first stage disk. Similarly, the interstage thermal seal is connected to the turbine disks through bolts in a circular pattern, typically clamping angular blade retaining rims to the opposite faces of the turbine disks as well. In addition, the second stage disk includes a rearwardly-extending cone which is bolted to the aft seal.
A disadvantage of such bolted connections is that they require holes to be formed in the disks which create stress concentrations and limit the useful lives of the seals and disks. Furthermore, additional disk weight is required to sustain the stresses imposed by the bolt and bolt hole engagements. Accordingly, recent turbine engines have been designed incorporating bayonet connections between the forward face plate, interstage seal, aft seal and the first and second stage disks.
A disadvantage with such bayonet connections is that alignment of the first and second stage disks and seals is difficult to maintain during operation, which may result in excessive vibrations during operation.
Attempts have been made to prevent rotation of the interstage seal with respect to the turbine disks by inserting an anti-rotation pin between the bore of the interstate seal and the first stage disk aft shaft. Disadvantages with this design is that the anti-rotation pin requires additional hardware such as a ring or wire for axial retention, and the anti-rotation torque carrying capability is reduced due to the pin being located at the bore.
Accordingly, there is a need for a device which minimizes alignment problems between the first and second stage disks and the interstage seal and specifically, there is a need for a design which eliminates rotation of the interstage seal in connection with the second stage disk that has increased torque carrying capability and does not require separate hardware for retention.